Scramjet combustor

ABSTRACT

A scramjet combustor for a hypersonic (Mach number greater than about 5.5) flight vehicle. The combustor housing has two spaced-apart, generally opposing, and longitudinally extending walls. Each wall has an aft-facing step. Fuel (such as hydrogen) is injected at the step into the shear zone along the separation line for better fuel-air mixing. The longitudinal distance between the steps is controlled to take advantage of aerothermo compression provided by the forward step&#39;s shock. The back portion of each wall is positioned transversely inward toward the combustor&#39;s longitudinal axis to improve mixing effectiveness for more efficient combustion.

This is a division of application Ser. No. 07/327,831, filed Sep. 26,1988.

BACKGROUND OF THE INVENTION

The present invention relates generally to a scramjet combustor for asupersonic flight vehicle and more particularly to a scramjet combustorhaving improved combustor efficiency and to a method for its operationwhich optimizes combustor performance.

Although the theory of scramjet engines has been well known for manyyears, and although supersonic combustors have been tested in thelaboratory, no scramjet engine is believed to have ever beensuccessfully flown. Recent advances in technology, such as in hightemperature materials, have made scramjet engines ready forimplementation in the next generation of high speed aircraft. Suchaircraft will be capable of flying at hypersonic speeds (i.e., speedshaving Mach numbers greater than about 5.5). Hypersonic flight vehicleshave been proposed which incorporate scramjet engines to achieve highMach numbers. Once such a vehicle has achieved a sufficient speed bysome other propulsive means (which may include a turbojet engine), ascramjet engine takes over to propel the aircraft to high Mach numbers(typically between Mach 6 and Mach 20). Such high Mach numbers cannot beachieved by any other known type of air-breathing engine.

A typical scramjet engine includes a combustor having a chamber, whereina fuel-air mixture moving at supersonic speed is burned, and having atleast one fuel injector which directs supersonically-moving fuel (suchas pressurized hydrogen) into the chamber. The engine also includes anair inlet, which delivers compressed supersonically-moving air to thecombustor chamber, and includes an exhaust nozzle, which channels theburning gases out of the combustor chamber to produce engine thrust. Thefuel injectors are the nozzle parts of the combustor to which fuel isdelivered by a fuel system which may include tanks, pumps, and conduits.

An important component of the scramjet engine is its combustor. Thebasic scramjet combustor of the literature includes alongitudinally-extending rectangular duct which defines the combustorchamber. The combustor's fuel injectors inject fuel into the combustorchamber through openings in the duct's two opposing larger walls. Thelongitudinally-moving air, from the engine inlet, and the typicallylongitudinally-or-transversely-injected fuel, from the fuel injectors,mix in the combustor chamber. In the case of hydrogen fuel, the fuel-airmixture in the combustor chamber will have a high enough temperature andpressure to auto-ignite.

The efficiency of burning within the combustor depends in part on howwell the air and fuel mix. To promote mixing, a scramjet combustordisclosed in the literature has included angled fuel injection whichmeans that the injected fuel is not parallel or perpendicular to thelongitudinally-moving air. Another approach disclosed in the literatureto promote better fuel-air mixing and burning stability has included anaft-facing step in one of the larger walls with (or without) theaddition of angled fuel injection at the step location. An additionalscramjet combustor, disclosed in the literature without elaboration, hasincluded an aft-facing step in each larger wall, with the steps being alongitudinal distance apart.

The efficiency of burning within the combustor also depends in part onhow much the air in the fuel-air mixture is compressed (increase instatic pressure). The more the air can be compressed (within the limitof temperature at which the air dissociates) before the fuel-air mixtureis burned, the more efficient and powerful the scramjet engine will be.The air compression disclosed in the literature has been accomplished bythe rectangular, funnel-like inlet portion of the engine which leadsfrom the engine inlet's entrance, where the engine inlet opening islargest, to the engine inlet's throat, where the engine inlet opening issmallest. The inlet of a scramjet engine may have fixed or variablegeometry. Variable geometry means that the engine throat area may bechanged, and there is an optimal throat area for any given set of flightconditions, as is known to those skilled in the art. A scramjet with avariable geometry engine inlet can operate, and operate moreefficiently, over a greater range of flight conditions than can ascramjet with a fixed geometry engine inlet. However, if the inletthroat area is made too small, air boundary layer instability or choking(reducing the airflow to sonic speed) in the inlet throat will exist,causing inlet unstart. This means that if the inlet compresses the airtoo much (because of a too small throat area) at lower Mach numbers, thescramjet engine cannot be started.

Although scramjet combustor designs have been proposed which increasecombustor efficiency, none are known which optimize combustor efficiencyby optimizing the parameters of step separation, fuel injection angle,and wall separation for design (cruise) flight conditions. Also, noscramjet combustors are known which are self-adaptive for suchparameters. This means none are known to have variable geometry for suchparameters enabling their configurations to be changed during supersonicflight to maintain optimized combustor efficiency during changes inflight conditions. Changes in flight conditions include, for example,changes in the combustor inlet Mach number during theacceleration-to-cruise phase of a flight.

SUMMARY OF THE INVENTION

It is an object of the invention to provide a scramjet combustor havingimproved burning efficiency for design flight conditions.

It is another object of the invention to provide a variable geometryscramjet combustor which can change its configuration during supersonicflight to maintain improved burning efficiency for off-design flightconditions.

It is a further object of the invention to provide a method foroperating such a variable geometry scramjet combustor to optimizeburning efficiency over a range of flight conditions.

There is provided by the invention described herein a scramjet combustorhaving first and second spaced-apart, generally opposing, and generallylongitudinally extending walls, with each of the walls having anaft-facing step with the step of the second wall spaced longitudinallyapart and aft of the aft-facing step of the first wall. The scramjetcombustor of the invention, optimized for the step separation parameter,has the longitudinal distance between the aft-facing steps generallybetween a minimum and a maximum value. At the minimum value, the shockfrom the step of the first wall, at the scramjet combustor's designinlet Mach number and fuel-air ratio, will strike the second wallproximate and longitudinally forward of the aft-facing step of thesecond wall. At the maximum value, the shock from the step of the firstwall, at the scramjet combustor's design inlet Mach number and fuel-airratio, will reflect off the second wall and then reflect off theseparation line emanating from the longitudinally forward step, as anexpansion fan whose initial expansion wave will strike the second wallproximate and longitudinally forward of the aft-facing step of thesecond wall. The separation line is a line or zone which separates thesupersonic air stream from a very low velocity air stream.

In another embodiment, the scramjet combustor of the invention hasvariable geometry for varying the step separation parameter whichincludes a mechanism for varying the longitudinal distance between theaft-facing steps during supersonic flight.

There is also provided a method for operating the scramjet combustor ofthe invention, with reference to the step separation parameter, whichincludes sensing a change in any of the input conditions of inlet Machnumber and fuel-air ratio and varying the longitudinal distance betweenthe aft-facing steps so as to keep generally between the above-describedminimum and maximum values.

In yet another embodiment of the invention, the scramjet combustor hastwo spaced-apart, generally opposing, and generally longitudinallyextending walls, with at least one of the walls having an aft-facingstep, the combustor also having a fuel injector positioned proximate thestep, wherein the scramjet combustor of the invention, optimized for thefuel injection angle, has the angle which the fuel injector makes withthe longitudinal axis, at the scramjet combustor's design inlet Machnumber and fuel-air ratio, generally equal to the angle which theseparation line makes with the longitudinal axis. Additionally, the fuelinjector positioned proximate the step includes a mechanism for varyingthe angle the fuel injector makes with the combustor's longitudinal axisduring supersonic flight. A method of the invention, with reference tothe fuel injection angle, includes sensing a change in any of the inputconditions of inlet Mach number and fuel-air ratio and varying the fuelinjection angle such that the angle which the fuel injector makes withthe longitudinal axis is kept generally equal to the angle which theseparation line makes with the longitudinal axis.

In a further embodiment of the invention, the scramjet combustor,includes a scramjet combustor housing having two spaced-apart, generallyopposing, and generally longitudinally extending walls. At least one ofthe walls has an aft-facing step, a generally longitudinally extendingfront portion, a transition portion, and a generally longitudinallyextending back portion. The front portion's forward end is attached tothe step's transversely outward end, the transition portion'stransversely outward terminus is attached to the front portion's aftend, and the back portion's front end is attached to the transitionportion's transversely inward terminus.

Yet another method of operating the scramjet combustor of the invention,with reference to the wall separation distance, includes makingmeasurements of inlet Mach number, fuel-air ratio, and inlet pressurelevel during supersonic flight and varying the transverse distancebetween the back portion of the combustor housing wall and the housing'slongitudinal axis during supersonic flight as a function of themeasurements so as to generally maintain a predetermined axialdistribution of static pressure and temperature within the combustor.

Several benefits and advantages are derived from the invention.Optimizing the longitudinal step separation distance will optimize shockcompression (from the shocks which form at the steps) for bettercombustion. Too short a distance will not allow a reflected shock (withits added aerothermo compression beyond the air compression limit of theengine inlet) to form from the longitudinally forward step's shock,while too long a distance will not add to the air compression, but doesadd to the combustor's longitudinal length and hence to its weight andfriction drag. Optimizing the fuel injection angle will promote betterfuel-air mixing allowing for more efficient combustion within a shorterlength combustor. Decreasing the transverse distance between the backportion of the combustor housing's walls will increase the pressurelevel and reduce the required fuel-air mixing distance, which willimprove combustion efficiency.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings illustrate a preferred embodiment of thepresent invention wherein:

FIG. 1 is a schematic cross-sectional view of a scramjet combustorincluding apparatus for varying step separation, fuel injection angle,and wall separation;

FIG. 2 is a schematic cross-sectional view showing the airflow insidethe combustor housing of FIG. 1 during supersonic flight;

FIG. 3 shows the airflow around the longitudinally forward step of FIG.1 at minimum step separation; and

FIG. 4 shows the airflow around the longitudinally forward step of FIG.1 at maximum step separation.

DETAILED DESCRIPTION OF THE INVENTION

A preferred scramjet combustor 10, shown in cross section in FIG. 1,includes a rectangular duct housing 12 forming a combustion chamber 14and having a forward air-inlet orifice 16 communicating with the engineinlet (not shown) and an aft air-outlet orifice 18 communicating withthe engine exhaust nozzle (also not shown). The combustor's longitudinalaxis 20 is defined by a line joining each orifice's center point (theorifice's "center of mass"). The housing 12 includes two spaced-apart,generally opposing, and generally longitudinally extending wider walls22 and 24 whose longitudinal edges are connected together by twonarrower walls to form a generally rectangular duct-shaped combustionchamber (only one 26 of the two narrower walls is shown in FIG. 1). Eachof such wider walls 22 and 24 includes an aft-facing step 28 and 30 witha transversely outward end 32 and 34, although some scramjet combustorapplications may require only a single step. Each wider wall 22 and 24also includes: a generally longitudinally extending front portion 36 and38 having a forward end attached to the step's transversely outward end32 and 34, a transition portion 40 and 42 having a transversely outwardterminus 44 and 46 attached to the front portion's 36 and 38 aft end,and a generally longitudinally extending back portion 48 and 50 having afront end attached to the transition portion's transversely inwardterminus 52 and 54 (such terminus extending towards the longitudinalaxis 20). The steps 28 and 30 are located a longitudinal distance 150apart, and the steps and the transition portions are not limited toplanar shape. The back portion 48 and 50 is disposed a transversedistance 152 from the longitudinal axis 20, with the distance from thetransition portion's transversely inward terminus 52 and 54 to thelongitudinal axis 20 being at least as great as the distance from thestep's transversely inward end 56 and 58 to the longitudinal axis 20.The transverse distance for each back portion is chosen such that thedesign (cruise conditions) inlet Mach number, fuel-air ratio, and inletpressure level would achieve a predetermined axial distribution ofstatic pressure and temperature within the combustor 10, as can be doneby those skilled in the art analytically using supersonic flowrelationships and equations and/or empirically through wind tunneland/or other laboratory testing. The required mixing distance is reducedby having the back portions 48 and 50 of the wider walls 22 and 24closer together.

Means are provided for adjusting the transverse distance of the largerwall's back portion 48 and 50 from the longitudinal axis 20 duringsupersonic flight. Preferably such means include hinged terminusattachments 60 for the transition portion 40 and 42 together with a pairof powered cylinders 62 to transversely move the back portion 48 and 50.The longitudinally spaced-apart powered cylinders 62 have their cylinderportions 64 retained by hinged attachments 66 fixed to a flight vehiclesupport structure 68 and their piston portions 70 retained by hingedattachments 72 fixed to the back portion 48 and 50. Alternate meansinclude powered cylinders each attached to both back portions outsidethe housing, as well as other positioning apparatus, as is known to theartisan. It is noted that a sliding seal arrangement could be used toallow the transverse movement of the wider walls 22 and 24 with respectto the narrower walls (such as 26).

Such means may be controlled by an onboard computer 74 which generatesan output signal 76 to vary the transverse distance of the back portion48 and 50 from the longitudinal axis 20 during supersonic flight. Inputs78 to the computer 74 would include measurements from a sensor array 80to provide inlet Mach number (defined to be the Mach number of the airat the combustor's air-inlet orifice 16), fuel-air ratio (defined to bethe ratio of the weight of fuel injected into the combustor's combustionchamber 14 per unit of time to the weight of the air entering thecombustor's air-inlet orifice 16 per unit of time), and inlet pressurelevel (defined to be the static pressure of the air at the combustor'sair-inlet orifice 16). The computer 74 would be programmed, as afunction of the measurements, to generally maintain a predeterminedaxial distribution of static pressure and temperature within thecombustor 10, as can be done by those skilled in the art. Suchprogramming would include the previously mentioned supersonic flowrelationships and equations and/or empirical data (which may be includedin a computer look-up table) from wind tunnel and/or other laboratorytesting. An alternate method of control would be, at some longitudinalpoint, to adjust the transverse distance of the back portion towards thelongitudinal axis until combustion occurred along the longitudinal axisor centerline (thereby achieving a desired static pressure andtemperature) as can be measured directly with an optical laserspectrometer set to detect the presence of water as a by-product ofcombustion, thereby detecting that combustion had taken place along thecenterline.

The combustor 10 also has a fuel injector 82 and 84 disposed proximateeach step 28 and 30 at an acute positive angle 154 (called the fuelinjector angle) with respect to the longitudinal axis 20 (supplementalfuel injectors 86 and 88 may be disposed along the front portion 36 and38 perpendicular to the longitudinal axis 20). During supersonic flight,a separation line 90 and 92 is produced by the step 28 and 30 which linevaries in accordance with the inlet Mach number and fuel-air ratio, asis known to those skilled in the art. On the transversely outward sideof the separation line 90 and 92, the air is relatively stagnant with arecirculation zone 94 and 96 near the step 28 and 30. On thetransversely inward side of the separation line 90 and 92, the air ismoving supersonically with a shear zone 98 and 100 (represented bycross-hatching in FIG. 2) near the separation line 90 and 92 itself. Thefuel injector angle is set to be generally equal to the angle which theseparation line makes with the longitudinal axis at the design (cruiseconditions) inlet Mach number and fuel-air ratio, as can be determinedby those skilled in the art analytically using supersonic flowrelationships and equations and/or empirically through wind tunneland/or other laboratory testing. By injecting the fuel 102 and 104 intothe shear zone 98 and 100 along the separation line 90 and 92, betterfuel-air mixing is achieved.

The scramjet combustor 10 further includes means for varying the fuelinjector angle during supersonic flight. Preferably such means includethe fuel injector's outlet nozzle 106 and 108 rotatably disposed in thestep 28 and 30 together with a powered cylinder 110 having its cylinderportion 112 fixed to a flight vehicle support structure 114 and itspiston portion 116 rotatably attached to the fuel injector's base 118and 120 through a connecting link 122 having hinged end attachments 124.Alternate means include the fuel injector's base having a pin disposedin a curved track with a powered cylinder for pin movement, as well asother angle positioning apparatus, as is known to the artisan.

Such means may be controlled by the onboard computer 74 which generatesan output signal 126 to vary the fuel injector angle during supersonicflight as a function of the step's separation line 90 and 92 formedduring supersonic flight for input conditions of inlet Mach number andfuel-air ratio. Inputs 78 to the computer 74 would include such measuredinput conditions from the sensor array 80. The computer 74 would beprogrammed, as a function of the measurements, to vary the fuelinjection angle during supersonic flight, when a change in any of theinput conditions is sensed, such that the fuel injector angle is keptgenerally equal to the angle which the separation line 90 and 92 makeswith the longitudinal axis 20, as can be done by those skilled in theart. Such programming would include the previously mentioned supersonicflow relationships and equations and/or empirical data (which may beincluded in a computer look-up table) from wind tunnel and/or otherlaboratory testing. An alternate method of control would be to directlymeasure the separation line angle optically from a shadowgraph, createdusing a high intensity light source, and then to adjust the fuelinjector angle to be equal to the measured value.

As seen from FIG. 2, during supersonic flight each step 28 and 30produces a shock 128 and 130, as well as a separation line 90 and 92,and such shock and separation line are a function of the inlet Machnumber and the fuel-air ratio, as is known to those skilled in the art.With the unfueled supersonic air flow 132 confined by the separationlines 90 and 92, with combustion essentially limited to the shear zone98 and 100 near the separation line 90 and 92, and with theclose-together back portions 48 and 50 of the wider walls 22 and 24pushing the separation lines 90 and 92 close together, the air flow isseen to be compressed and combustion is seen to occur transverselyacross essentially the entire air stream. This desirable result givesmore efficient combustion and a shorter combustor length, theimprovement being unattainable in the prior art.

To optimize shock compression, the longitudinal distance between the twosteps 28 and 30 is chosen to be between a minimum and a maximum value.At the minimum value, for a design inlet Mach number and fuel-air ratio,the shock 128 of the longitudinally forward step 28 of the correspondingfirst 22 of the wider walls will strike the second of the wider walls 24proximate and longitudinally forward of that wall's longitudinally aftstep 30 as shown in FIG. 3. At the maximum value, for a design inletMach number and fuel-air ratio, the shock 128 of longitudinally forwardstep 28 will reflect off the second of the wider walls 24 and thenreflect off the longitudinally forward step's separation line 90 as anexpansion fan whose initial expansion wave 134 will strike the second 24of the wider walls proximate and longitudinally forward of the aft step30 as shown in FIG. 4. The minimum and maximum values for the design(cruise conditions) inlet Mach number and fuel-air ratio can bedetermined by those skilled in the art analytically using supersonicflow relationships and equations and/or empirically through wind tunneland/or other laboratory testing. Preferably, the longitudinal distanceis chosen to be generally equal to the minimum value. As the air flowcrosses a shock, it is compressed because its Mach number decreases andits static pressure increases.

The scramjet combustor 10 additionally includes means for varying thelongitudinal distance 150 between the steps 28 and 30 during supersonicflight. Preferably such means include the longitudinally aft step's 30larger wall 24 having a longitudinally overlapping section with an innerwall portion 136 extending longitudinally forward toward the combustor'sair-inlet orifice 16 and with an outer wall portion 138 extendinglongitudinally aft toward the aft step 30. Such means also include apowered cylinder 140 for that larger wall's 24 front portion 38. Thegenerally longitudinally disposed powered cylinder 140 has its cylinderportion 142 fixed to a flight vehicle support structure 144 and itspiston portion 146 fixed to the front portion 38 of the aft step's 30larger wall 24. Alternate means include a rack and pinion arrangementfor the front portion, as well as other positioning apparatus, as isknown to the artisan.

Such means may be controlled by the onboard computer 74 which generatesan output signal 148 to vary the longitudinal distance between the steps28 and 30 during supersonic flight as a function of input conditions ofinlet Mach number and fuel-air ratio. Inputs 78 to the computer 74 wouldinclude such measured input conditions from the sensor array 80. Thecomputer 74 would be programmed, as a function of the measurements, tovary the longitudinal distance between the steps 28 and 30 duringsupersonic flight, when a change in any of the input conditions issensed, such that the longitudinal step separation distance would begenerally between the previously described minimum and maximum values(and preferably be generally equal to the minimum value). Suchprogramming would include the previously mentioned supersonic flowrelationships and equations and/or empirical data (which may be includedin a computer look-up table) from wind tunnel and/or other laboratorytesting. An alternate method of control would be to directly opticallymeasure the point where the longitudinally forward step's shock strikesthe second of the larger walls from a shadowgraph, created using a highintensity light source, and then to adjust the step separation so thatthe strike point falls within those strike points corresponding to theminimum and maximum step separation values.

It is noted that the invention provides a more efficient scramjetcombustor, one which utilizes variable step separation distance forseries type aerothermo (shock) compression, variable fuel injectionangle for improved fuel-air mixing, and variable transverse wallseparation for reduced fuel-air mixing length.

The foregoing description of a preferred embodiment of the invention hasbeen presented for purposes of illustration. It is not intended to beexhaustive or to limit the invention in the precise form disclosed (suchas, for example, a particular number of steps or a particular shape fora step or a transition portion), and obviously many modifications andvariations are possible in light of the above teaching.

I claim:
 1. A method for operating a scramjet combustor, said combustorincluding a housing having a forward air-inlet orifice and an aftair-outlet orifice each with a center point together defining alongitudinal axis, said housing also having two spaced-apart, generallyopposing, and generally longitudinally extending walls, wherein at leastone of said walls includes: an aft-facing step with a transverselyoutward end; a generally longitudinally extending front portion having aforward end attached to said transversely outward end of said step, andhaving an aft end; a transition portion having a transversely outwardterminus attached to said aft end of said portion and having atransversely inward terminus extending towards said longitudinal axis;and a generally longitudinally extending back portion having a front endattached to said transversely inward terminus of said transitionportion; said back portion disposed a transverse distance from saidlongitudinal axis, said method comprising:(a) making measurements ofinput conditions of inlet Mach number, fuel-air ratio, and inletpressure level during supersonic flight; and (b) varying said transversedistance during supersonic flight as a function of said measurements togenerally maintain a predetermined axial distribution of static pressureand temperature within said combustor.